1. Field of the Invention
The present invention relates to improvements in a search and tracking radar system which is useful for the navigation of a spacecraft, for example, between a plurality of space stations.
2. Description of the Prior Art
First of all, a typical application of a radar system of this category will be described in terms of the conception of working on an assumption that the radar system is used as a navigation radar for the navigation of a spacecraft between a plurality of space stations. Referring to FIG. 1 showing the conception of working in the above-mentioned application of a radar system, there are shown a spacecraft 1, space stations 2a and 2b, the Earth 3, a first radio wave A1 transmitted from the spacecraft 1 and second radio waves A2 and A3 transmitted from the space stations 2a and 2b respectively. The second radio wave is generated in two cases, namely, in a case in which the first radio wave is reflected by the space station and in a case in which the space station generates internally the second radio wave.
Now, suppose that the spacecraft 1 approaches the space station 2a or 2b for docking. A navigation radar put aboard the spacecraft 1 makes a search for the space station which is a desired target. After the target has been detected, the navigation radar functions as a sensor for guiding the spacecraft 1. The objective space of the search is the entire space observable from the spacecraft 1. Accordingly, it is necessary to radiate radio waves simultaneously in all directions from the spacecraft 1 or to scan the entire space with a narrow beam.
Referring to FIG. 2 showing the constitution of a conventional radar system, there are shown a monopulse antenna 4, a sum antenna radiation pattern 5a, a difference antenna radiation pattern 5b in the direction of elevation, a difference antenna radiation pattern 5c in the azimuthal direction, a transmitter 6, a duplexer 7, a comparator 8, a receiver 9, a range tracker 10, an AGC (Automatic Gain Controller) 11, a goniometer 12, an antenna driving unit 13, a transponder antenna 14 and a transponder 15. The elements designated by reference characters 4 through 13 are put aboard the spacecraft 1 and the antenna 14 and the transponder 15 are placed on the space station 2a.
The first radio wave A1 generated by the transmitter 6 is radiated through the duplexer 7 and the monopulse antenna 4 toward the space station 2a. The monopulse antenna 4 consists of four horn antennas. These four horn antennas are used simultaneously to transmit the radio wave via the sum antenna radiation pattern 5a. The transponder antenna 14 receives the first radio wave A1 and supplies the same to the transponder 15. In response to the reception of the first radio wave A1, the transponder 15 generates after a delay of a fixed time the second radio wave A2, which is radiated into the external space through the transponder antenna 14. The four horn antennas constituting the monopulse antenna 4 receive the second radio wave and supply the same through the duplexer 7 to the comparator 8. The comparator 8 processes the four-channel input signals through addition or subtraction and gives known signals .SIGMA., .DELTA..sub.EL and .DELTA..sub.AZ of three channels to the receiver 9. The .SIGMA.-signal corresponds to the signal given by receiving the second radio wave A2 by the antenna radiation pattern 5a, while the .DELTA..sub.EL -signal and the .DELTA..sub.AZ -signal correspond to the signals given by receiving the second radio wave A2 by the difference antenna radiation pattern 5b in the direction of elevation and the difference antenna radiation pattern 5c in the azimuthal direction respectively. The receiver 9 amplifies and detects the input signal of each channel and gives output signals to the goniometer 12. Part of the .SIGMA.-signal among the output signals of the receiver 9 is transferred to the range tracker 10. The range tracker 10 decides the time of reception of the second radio wave A2 and obtains the distance R from the spacecraft 1 to the space station 2a by the use of Equation (1). ##EQU1## where t.sub.E is the time when the second radio wave A2 is received, t.sub.S is the time when the first radio wave A1 is transmitted, t.sub.D is the delay time before the radio wave generation by the transponder 15 and C is the velocity of light. The time t.sub.S is transferred from the transmitter 6 to the range tracker 10. t.sub.D is a known quantity.
The range tracker 10 generates gate pulses to actuate the receiver 9 only while the second radio wave A2 is received. The AGC 11 receives the gate pulses and controls the gain of the receiver 9 so that the gain of the receiver 9 is zero except while the voltage of the gate pulses is higher than a value. Consequently, the signals of the .SIGMA.-channel, .DELTA..sub.EL -channel and .DELTA..sub.AZ -channel are given to the goniometer 12 only while the second radio wave A2 is received. The goniometer 12 processes those input signals of the three channels and determines the direction of arrival of the second radio wave A2 according to the well-known monopulse goniometry employed in the monopulse radar. The angle information (.theta..sub.0, .phi..sub.0) is transferred together with distance information R given by the range tracker 10 to orbit computing means. The angle information and the distance information are used for deciding and correcting the docking orbit.
The angle information given by the goniometer 12 is transferred also to the antenna driving unit 13. The antenna driving unit 13 scans the monopulse antenna 4 on the basis of the angle information and drives the antenna so that the second radio wave A2 is received in alignment with the direction of the center line of the antenna to track the space station 2a.
In a situation as shown in FIG. 1, the spacecraft 1 does not always have the definite information of the direction of the space station 2a, i.e., the target of docking, and in some cases, it is possible that the spacecraft 1 is obliged to change the docking target from the space station 2a to the space station 2b during the approach to the space station 2a. Then, the radar system put aboard the spacecraft 1 needs first to make a search to catch the space station 2a or 2b in the field of view of the radar system. As well known, a monopulse radar uses a high gain antenna and the beam width of such an antenna usually is 10.degree. or less. Accordingly, the above-mentioned search requires the scanning of antennas. Differing from the scanning operation of most terrestrial radar systems, the scanning of the entire space by the radar put aboard a spacecraft is not easy and requires much time. In addition, the operation of the mechanical antenna driving unit 13 for the search for targets affects adversely the payload of the spacecraft 1 or, even in the case of searching space stations through the attitude control of the spacecraft 1 itself, the fuel consumption of the spacecraft 1 is increased significantly. Furthermore, even after the space station has been caught and the tracking operation has been started, the space station needs to be always caught on the line of sight of the monopulse antenna for highly accurate tracking, and the highly accurate direction control of the antenna is necessary to make the spacecraft 1 approach the space station by means of an advanced navigation such as the proportional navigation. The use of a phased array antenna, which electrically scans the beam, instead of a mechanically scanned antenna eliminates mechanical means, however, since the phased array antenna forms a narrow beam, the phased array antenna is not suitable, similarly to a mechanically scanned antenna, for scanning the entire space and hence the phased array antenna can not reduce the time required to search for the space station.